The present invention relates to a separator for an annular gas turbine engine combustion chamber which separates the combustion zones of such a combustion chamber having two or more combustion zones within the combustion chamber.
In order to achieve lower pollution, modem gas turbine aircraft engines make use of an annular combustion chamber having two or more arrays of fuel injectors injecting fuel into two or more combustion zones within the combustion chamber. One of the fuel injector arrays and one of the combustion zones operates during low power engine operation, while the other fuel injector array and combustion zone operates under take-off or full power conditions.
Typically, such known combustion chambers comprise two mutually spaced apart, generally annular walls extending around a longitudinal axis of the engine joined at an upstream end by a wall which extends generally radially relative to the longitudinal axis and interconnects the spaced apart annular walls. Two distinct arrays of fuel injectors, both of which extend generally in an annular array about the longitudinal axis extend through the upstream end wall of the combustion chamber and serve to inject fuel into two combustion zones within the combustion chamber. The first array of fuel injectors supplies fuel in a first mode of engine operation, such as under low power operating conditions, while the second array of fuel injectors supplies fuel into a second combustion zone during a second mode of engine operation, such as at take-off under full-power. Oxidizer intake passages may also be defined by the upstream end wall to enable oxidizer to pass into the combustion chamber zones so as to support combustion therein. It is also known to have a gas separator extending from the upstream end wall of the combustion chamber into the combustion chamber located between the two annular arrays of fuel injectors to separate the first and second combustion zones.
The gas flow separator assembly represents a critical factor in designing a combustion chamber since it is subjected to longitudinal and transverse stresses during the combustion chamber operation. In a typical gas turbine engine of this type, ignition of the fuel initially takes place in a first combustion zone which is served by the pilot injectors during lower power operating conditions. Under these conditions, typically while the aircraft is on the ground, combustion is stabilized within the first combustion zone and is supplied fuel only by the first array of pilot injectors, which are the only injectors operating. During engine acceleration, typically an operation corresponding to approximately 20% of the take-off thrust, the second fuel injector array, or take-off injectors supply fuel to the second combustion zone. The fuel and oxidizer mixture in the second combustion zone is ignited as the flame propagates from the first combustion zone served by the pilot injectors towards the second combustion zone. It is quite important that fuel from all of the second injector array be ignited substantially simultaneously to insure proper engine operation.
However, ignition malfunctions of the take-off fuel injector array have been repeatedly observed. Such malfunctioning is linked to the lack of direct propagation of the pilot injector flame toward the second combustion zone served by the take-off injectors, the flame front being constrained to move around the gas flow separator between the two arrays of fuel injectors.
European Patent Application 0 564 170 describes a gas flow separator assembly comprising a plurality of distinct separator segments located adjacent to each other to form an annular separator in which each of the separate separator segments are independent of adjacent separator segments. Each of the separator segments comprises an elongated body having opposite sides which are spaced from corresponding sides of adjacent segments.